This invention relates to spacecraft attitude control systems and, more particularly, to such attitude control systems using reaction wheel assemblies.
Spacecraft control may relate to the location of the spacecraft, also known as stationkeeping in the case of a geosynchronous spacecraft, and also includes attitude control, which establishes the orientation of the various faces of the spacecraft relative to some other body. Control of the location of the spacecraft requires accelerating the spacecraft as a whole in some desired direction, which is performed by ejecting matter in the opposite direction by the use of thrusters. Attitude control involves the application of torques to the spacecraft for proper orientation.
Torques may be applied to a spacecraft by ejecting matter from thrusters at diverse locations on the spacecraft so as to create torques. This may be accomplished, for example, by propellant or monopropellant chemical thrusters, with or without electrical augmentation. Such thrusters have the major disadvantage that the supply of ejectable matter or propellant may become depleted, thereby terminating the ability to perform attitude control maneuvers. Chemical thrusters also have the disadvantage that they are difficult to modulate to perform fine attitude control maneuvers. Another method for torquing spacecraft for attitude control involves the use of magnetic coils associated with the spacecraft, which when energized create magnetic fields which interact with the geomagnetic field of the earth to generate torques. Magnetic torquers have the advantage of being energized by electricity, which on a spacecraft equipped with solar panels is a renewable resource. However, magnetic torquing has the disadvantage that the torquing is dependent not only on the current flowing through the coils, but also upon the earth's magnetic field, which varies from place to place and from time to time, which makes accurate control difficult. Also, magnetic torquing provides relatively modest levels of torque which are not suitable for rapid slewing, such as may be required for repositioning antennas or sensors, or for recovering from transient disturbances.
Torquing of the spacecraft may also be provided by reaction wheels or momentum wheels, the inertia of which reacts to provide body torque when the wheel is accelerated. In general, "reaction" wheels provide torques by accelerations which cause the wheel to spin in either direction, passing through zero angular velocity at some times, whereas momentum wheels are "biased" to a relatively high angular velocity, and are accelerated and decelerated, as required, while maintaining rotation in a single direction. Such wheels have the advantages of precise control and electrical operation. In order to have significant torquing effect, the wheel must have significant mass, significant size, or be capable of high angular velocities. Because of weight constraints on the spacecraft, it is desirable to minimize the mass of the wheel. Reductions in the mass or in the wheel diameter decrease the inertia, and increase the angular velocities required for proper spacecraft torquing. Compromises must be made among the various parameters, depending upon the spacecraft mission. As between momentum wheels and reaction wheels, the reaction wheel is subject to attitude errors due to bearing friction effects which occur at and near zero angular velocity, as described in U.S. Pat. No. 5,020,745, issued Jun. 4, 1991 in the name of Stetson, Jr. Various schemes have been devised for correcting for the effects of friction in a reaction wheel, as for example that described in copending application, Ser. No. 07/732,963, filed Jul. 19, 1991 in the name of Goodzeit et al, now U.S. Pat. No. 5,201,833. The momentum wheel has the advantage that the wheel bearings tend to be more reliable than the bearings in a reaction wheel, which is believed to result from maintenance of a continuous film of lubricant on the bearing surfaces by the rapid motion of the bearing. The momentum wheel also avoids the problem of low speed bearing friction by biasing the wheel speed so that the wheel never reaches zero angular velocity. However, this reduces the potential range of control by comparison with that of the reaction, because the reaction wheel may be accelerated to both positive and negative angular velocities, whereas the momentum wheel cannot have both positive and negative velocities.
When reaction or momentum wheels are used for attitude control, it is found that the control requirements may result in a continuous average secular increase or decrease in the velocity of a wheel. This is most easily understood by considering a geosynchronous, earth-facing communications satellite, which must make one complete rotation every 24 hours, as its orbit takes it around the earth, in order to maintain a face directed toward earth. The torque required to provide this slow continuous rotation of the spacecraft body is provided by a continuous acceleration or deceleration of a wheel whose axis is perpendicular to the orbital plane. The wheel speed cannot continue to increase indefinitely, because centrifugal effects would result in disintegration. In the case of a momentum wheel, its speed cannot be permitted to decrease to zero angular velocity. As a result, a procedure known as "unloading" of the wheel is applied by which thrusters or magnetic torquers are energized for relatively short periods to generate torques which, when applied to the spacecraft body, allow the wheel speed to be returned to a nominal value. Such a scheme using magnetic torquers is described, for example, in allowed application Ser. No. 07/488,919, filed Mar. 5, 1990 in the name of Linder et al, now U.S. Pat. No. 5,123,617.
The very large capital cost of satellites and of placing satellites into orbit requires that the expected lifetime of the spacecraft be extended as much as possible so as to achieve a suitable economic return during its operating life. For this purpose, the spacecraft mass is minimized, so the maximum amount of expendable propellant for the thrusters can be carried into orbit, and various redundancy schemes are provided, both in the spacecraft and in its payload, to insure that a single failure does not result in failure of the mission. Among these redundancy schemes is the use of a number of reaction or momentum wheels which exceeds the minimum number required for control of the spacecraft attitude. For example, in a three-axis stabilized spacecraft, the minimum number of wheels required for three-axis control is three, one along each axis of control. However, to take into account the possibility of failure of a single wheel, four or more wheels with mutually skewed axes may be provided. While such a scheme provides redundancy, it substantially increases the complexity of the control system, in that controlled acceleration of any one wheel results in torques of the spacecraft body about two or more principal axes. Also, when four or more wheels are used for three-axis control, the amount of electrical energy applied to the wheels may exceed the minimum required for control. A scheme for controlling the wheel speeds in a system including four or more wheels for minimizing the power is described in U.S. Pat. No. 5,058,835, issued Oct. 22, 1991 in the name of Goodzeit et al.